This invention relates to a vane of a gas turbine engine and more specifically to a configuration that provides improved heat transfer to the trailing edge region of the vane while minimizing pressure loss to the cooling fluid.
As performance requirements for gas turbine engines increase, operating temperatures increase as well, especially in the turbine section. While technological advancements have been made in many areas including material capability and thermal barrier coating systems to withstand these higher operating temperatures, it is also desirable to obtain as efficient cooling of the component as possible. In a gas turbine engine, the turbine section is comprised of alternating rows or stages of vanes and blades, where the vanes remain stationary and the blades rotate about the engine axis. The vanes serve to direct the flow of hot gases to the next stage aft of a turbine, onto a set of rotating blades. The orientation at which this flow of hot gases is directed is critically important to the overall turbine performance and blade life. Therefore, it is necessary to ensure that the vane trailing edge shape is maintained and proper cooling of the vane trailing edge is one means to accomplish this objective.
Prior art turbine vanes have incorporated pedestals or pin fins in the vane walls to aid in cooling and heat transfer by causing turbulation in the wake regions generated by cooling fluid passing around the pedestals or pin fins. These pedestals are often times located towards the vane trailing edge. A prior art example of vane trailing edge cooling utilizing pin fins is disclosed in U.S. Pat. No. 4,515,523 where pin fins are added to the rib walls that extend longitudinally to the trailing edge for increased stiffness. These additional pin fins serve to replace those eliminated due to the placement of longitudinal ribs. However, the placement of these additional pin fins along the rib wall causes additional pressure loss to the cooling flow. The present invention seeks to overcome the shortcomings of the prior art by providing a vane trailing edge region having the required stiffness through longitudinal ribs and improved heat transfer associated with pin fins while reducing the pressure loss to the cooling flow. The reduced pressure loss along the rib wall is a result of repositioning the pedestals closer to the rib walls in conjunction with incorporating recessed cavities in the rib walls in areas immediately adjacent to the pedestals.
The present invention provides a gas turbine vane having first and second platforms in spaced relation, an airfoil extending between the platforms, with the airfoil containing one or more cooling circuits. The cooling circuit has a row of first pedestals having a first diameter, one or more rows of second pedestals having a second diameter and spaced a first distance axially from the first pedestals and offset radially a second distance, and one or more rows of third pedestals having a third diameter and spaced a third distance axially from the second pedestals and offset radially a fourth distance. A plurality of generally axially extending ribs are incorporated with the ribs bisecting the rows of first, second, and third pedestals, and with the ribs having at least one recessed cavity in each of its upper and lower walls. The recessed cavities are positioned immediately adjacent second and third pedestals located closest to the ribs such that a cavity passageway is formed to pass sufficient cooling fluid between the rib and pedestal. The recessed cavities allow for closer positioning of pedestals to the rib to enhance the overall heat transfer while minimizing pressure loss.
It is an object of the present invention to provide a vane for a gas turbine engine having improved heat transfer and reduced pressure loss to the cooling fluid.
In accordance with these and other objects, which will become apparent hereinafter, the instant invention will now be described with particular reference to the accompanying drawings.